1. Field of the Invention
The present invention generally relates to the art of rocket propulsion systems, and more specifically to an improved configuration for a hybrid rocket engine which utilizes a tank filled with high pressure non-flammable gas to pressurize a liquid propellant component container.
2. Description of the Related Art
A hybrid rocket engine or motor is a cross between a solid propellant rocket and a liquid propellant rocket. A hybrid rocket utilizes a liquid oxidizer to burn a solid fuel element. A reverse hybrid rocket applies a combustible liquid fuel to a solid oxidizer. The hybrid rocket propellant can be ignited by an igniter such as an electrically generated spark, or by initial injection of an ignition fluid which exothermically reacts with the liquid oxidizer. An example of a conventional hybrid rocket configuration is presented in U.S. Pat. No. 3,323,308, issued Apr. 9, 1964, entitled "CONSTANT FLOW, VARIABLE AREA HYBRID ENGINE INJECTOR", to J. Greco.
A hybrid engine provides the following basic advantages over a purely solid or liquid fuel rocket engine: (1) the complete separation of fuel from the principal oxidizer, eliminating the potential for uncontrolled mixing, (2) the capability to use an optimum combination of propellant ingredients regardless of whether these are solid or liquid, and (3) the capability to easily stop and restart the engine. Hybrid engines have the potential for low cost, reliability, and safety. In addition to its on-off capability, the engine is easily throttleable since there is only one liquid component. Since the solid fuel components need not contain any oxidizer, they are easily produced under less hazardous conditions.
A conventional hybrid rocket engine includes a hollow housing or combustion chamber in which an elongated solid fuel propellant component or grain is mounted. The fuel grain typically has a "wagon wheel" cross section, with a central hollow hub, a rim, and a plurality of spokes joining the hub to the rim. The spaces between the spokes are hollow, allowing flow of combustion gas through the length of the fuel grain. Liquid oxidizer is provided in a tank or container mounted forward of the fuel grain, and caused to flow through the fuel grain and out a nozzle mounted at the rear end of the combustion chamber. Ignition causes combustion of the fuel-oxidizer mixture at the surfaces of the fuel grain, resulting in the generation of thrust as the high pressure combustion products are discharged out of the combustion chamber through the nozzle.
It is essential to sustain the flow of the liquid propellant component into the combustion chamber for as long a period of time and with as constant a pressure as possible to completely utilize the fuel grain. Methods of pressurizing the liquid component container to cause the liquid flow therefrom at a controlled rate include: (1) stored gas systems, (2) oxidizer evaporation systems, (3) systems evaporating non-propellants, (4) systems using products of chemical reactions, (5) autogenous systems and auxiliary pumping systems.
The present invention improves on the type (1) system which utilizes a stored gas, or alternatively a turbo-pump feed system. In the conventional arrangement, a non-flammable gas such as helium or nitrogen is provided in a high pressure cylinder or tank which is mounted external of the combustion chamber and liquid propellant component container. The high pressure gas is fed into the container through an appropriate regulator or throttle valve to create a pressure in the container selected to force the liquid propellant component into the combustion chamber at a desired flow rate.
A major problem inherent in the conventional hybrid rocket engine design is that it is difficult to burn all of the fuel grain in a controlled manner. Where an attempt is made to burn the entire fuel grain, a point is reached where a central unburned portion separates from the main body from the grain, and is moved by the combustion gas flow into the nozzle area, causing a major reduction in thrust. This catastrophic change may be avoided by burning the fuel grain only down to a point at which separation will not occur. However, this results in wasting the remaining portion of the fuel grain, and thereby precious space and weight in the rocket motor.
The wagon wheel grain configuration represents an optimal ratio of exposed surface area to cross sectional area. However, the central hub is generally hollow, and thereby unused, representing wasted space in the motor. Since the hub is supported only by the spokes which become progressively burned away during operation of the motor, the possibility that the hub will deleteriously separate from the main body of the grain toward the end of the fuel burn as discussed above is substantial.
Another problem inherent in a conventional stored gas pressurization system is that the pressurizing gas is stored in the tank at very high pressure, on the order of 6000 psi. Structural failure of the tank would create an extremely dangerous situation, possibly resulting in serious injury to personnel and damage to facilities in the vicinity of the ruptured tank.